Gas turbine engine blade outer air seal profile

ABSTRACT

A blade outer air seal for a gas turbine engine includes a gas path surface exposed to exhaust gas flow, a first side extending radially outward from the gas path surface, a second side extending radially outward from the gas path surface, and a plurality of film cooling holes disposed on at least one of the gas path surface. The first side and the second side, the film cooling holes are disposed at locations described by a set of Cartesian coordinates set forth in Table 1. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate and a radial coordinate relative to a defined point of origin. A gas turbine engine is also disclosed.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-energy exhaust gas flow. The high-energy exhaust gas flow expandsthrough the turbine section to drive the compressor and the fan section.The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

Both the compressor and turbine sections include rotating bladesalternating between stationary vanes. The vanes and rotating blades inthe turbine section extend into the flow path of the high-energy exhaustgas flow. Leakage around vanes and blades reduces efficiency of theturbine section. Blade outer air seals control leakage of gas flow andimprove engine efficiency. All structures within the exhaust gas flowpath are exposed to the extreme temperatures. A cooling air flow istherefore utilized over some structures to improve durability andperformance.

SUMMARY

In a featured embodiment, a blade outer air seal for a gas turbineengine includes a gas path surface exposed to exhaust gas flow, a firstside extending radially outward from the gas path surface, a second sideextending radially outward from the gas path surface, and a plurality offilm cooling holes disposed on at least one of the gas path surface. Thefirst side and the second side, the film cooling holes are disposed atlocations described by a set of Cartesian coordinates set forth inTable 1. The Cartesian coordinates are provided by an axial coordinate,a circumferential coordinate and a radial coordinate relative to adefined point of origin.

In another embodiment according to the previous embodiment, includes anaxially forward side and an axially aft side. The gas path surface andthe forward side define an arc and the point of origin is defined at thecenter of curvature of the arc on the forward side.

In another embodiment according to any of the previous embodiments, theforward side and the aft side include features for securement to asupport structure within the turbine section of the gas turbine engine.

In another embodiment according to any of the previous embodiments, theblade outer air seal is one of a plurality of outer air seals disposedcircumferentially about a longitudinal axis of the gas turbine engine.

In another embodiment according to any of the previous embodiments, eachof the film cooling air holes are located within a true position of0.023 inches (0.58 mm).

In another embodiment according to any of the previous embodiments, atleast some of the film cooling air holes include one of a conical andcylindrical shape.

In another embodiment according to any of the previous embodiments, eachof the film cooling air holes correspond with a passage through thecorresponding surface and at least some of the passages are disposed atan angle different than normal relative to the surface.

In another embodiment according to any of the previous embodiments, aplurality of film cooling holes have a diameter within a range of0.010-0.035 inch (0.25-0.89 mm).

In another featured embodiment, a gas turbine engine includes acompressor section disposed about an axis. A combustor is in fluidcommunication with the compressor section. A turbine section is in fluidcommunication with the combustor. The turbine section includes at leastone rotor having a plurality of rotating blades. A plurality of bladeouter air seals circumferentially surround the rotating blades. Each ofthe blade outer air seals includes a gas path surface exposed to exhaustgas flow, a first side extending radially outward from the gas pathsurface, a second side extending radially outward from the gas pathsurface, and a plurality of film cooling holes disposed on at least oneof the gas path surface. The first side and the second side, the filmcooling holes are disposed at locations described by a set of Cartesiancoordinates set forth in Table 1. The Cartesian coordinates provided byan axial coordinate, a circumferential coordinate and a radialcoordinate relative to a zero-coordinate.

In another embodiment according to any of the previous embodiments, eachof the blade outer air seals includes an axially forward side and anaxially aft side. The gas path surface and the forward side define anarc and the zero-coordinate is defined at the center of curvature of thearc on the forward side.

In another embodiment according to any of the previous embodiments, theforward side and the aft side include features for securement to asupport structure within the turbine section of the gas turbine engine.

In another embodiment according to any of the previous embodiments, eachof the plurality of film cooling air holes are located within a trueposition of 0.023 inches (0.58 mm).

In another embodiment according to any of the previous embodiments, atleast some of the plurality of film cooling air holes include one of aconical and cylindrical shape.

In another embodiment according to any of the previous embodiments, eachof the plurality of film cooling air holes are in communication with acorresponding plurality of passages and at least some of the passagesare disposed at an angle different than normal relative to the surface.

In another embodiment according to any of the previous embodiments, aplurality of film cooling holes have a diameter within a range of0.010-0.035 inch (0.25-0.89 mm).

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is a schematic view of a portion of a turbine section of a gasturbine engine.

FIG. 3 is a perspective view of an example blade outer air seal.

FIG. 4 is a another perspective view of the example blade outer airseal.

FIG. 5 is a first side view of an example blade outer air seal.

FIG. 6 is a second side view of an example blade outer air seal.

FIG. 7 is a schematic view through a film cooling hole.

FIG. 8A is a schematic view of a film cooling hole.

FIG. 8B is a cross-section schematic view of the film cooling hole ofFIG. 8A.

FIG. 9 is a schematic view of an example film cooling hole.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 58 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 58 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 58 includes airfoils 60 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10.67 km). The flight condition of 0.8 Mach and35,000 ft (10.67 km), with the engine at its best fuel consumption—alsoknown as “bucket cruise Thrust Specific Fuel Consumption ('TSFC)”—is theindustry standard parameter of lbm of fuel being burned divided by lbfof thrust the engine produces at that minimum point. “Low fan pressureratio” is the pressure ratio across the fan blade alone, without a FanExit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosedherein according to one non-limiting embodiment is less than about 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram°R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second (350 m/second).

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

Referring to FIG. 2, the example turbine section 28 includes at leastone rotor 34 having a turbine blade 62. The turbine blade 62 includes atip 65 disposed adjacent to a blade outer air seal 70 (BOAS). Astationary vane 66 is mounted and supported within the case 64 on atleast one side of the turbine blade 62 for directing gas flow into thenext turbine stage. The BOAS 70 is disposed adjacent to the tip 65 toprovide a desired clearance between the tip 65 and a gas path surface 72of the BOAS 70. The clearance provides for increase efficiency withregard to the extraction of energy from the high energy gas flowindicated by arrow 68.

The turbine blade 62 and vane 66 along with the blade outer air seal areexposed to the high-energy exhaust gas flow 68. The high energy exhaustgas flow 68 is at an elevated temperature and thereby structures such asthe blade 62, vane 66 and the BOAS 70 are fabricated from materialscapable of withstanding the extremes in temperature. Moreover, each ofthese structures may include provisions for generating a cooling filmair flow 75 over the surfaces. The cooling film air flow generates aboundary layer that aids in survivability for the various structureswithin the path of the exhaust gasses 68.

In the discloses example, a plurality of BOAS 70 are supported withinthe case 64 and abut each other to form a circumferential boundaryradially outward of the tip 65. Accordingly, at least one stage of theturbine section 28 includes a plurality of BOAS 70 that define a radialclearance between the tip 65 and the gas path surface 72. Additionalstages in the turbine section 28 will include additional BOAS to definethe radial clearance with turbine blades of each stage.

Referring to FIGS. 3 and 4 with continued reference to FIG. 2, the BOAS70 includes a plurality of film cooling holes 86 for generating a filmcooling air flow, indicated at 75 in FIG. 2, along the gas path surface72. The film cooling holes 86 are disposed on surfaces exposed to theexhaust gasses 68. It should be understood that the term “holes” is usedby way of description and not intended to limit the shape to a roundopening. Accordingly, the example holes 86 maybe round, oval, square orany other shape desired.

Referring to FIGS. 5 and 6 with continued reference to FIGS. 3 and 4,the example BOAS 70 includes the gas path surface 72 that is exposeddirectly to the exhaust gasses 68. The BOAS 70 further includes a firstside 74 and a second side 76. The first and second sides 74, 76 abutadjacent BOASs disposed circumferentially about the turbine case 64.Each of the BOASs 70 includes a forward surface 78 and an aft surface80. The forward surface 78 and aft surface 80 includes support featuresfor holding each BOAS within the turbine case 64. In this example, theBOAS 70 includes a forward channel 90 and an aft tab 88 to conform tofeatures within the turbine cases 64 to support the BOAS 70circumferentially about the corresponding turbine blade 62.

The first side 74 and second side 76 and a gas path surface 72 allinclude a plurality of film cooling holes 86. Each of the film coolingholes 86 provide a pathway for cooling air to generate the boundarylayer of cooling air flow 75 to maintain the BOAS within definedtemperature ranges. A specific location of the film cooling holes 86 isdevised to provide cooling air flow coverage of features susceptible tothe high temperature exhaust gasses. The cooling holes 86 are arrangedto produce boundary layers of cooling flow along the gas path surface 72along with the first side 74 and the second side 76. As appreciated, thefirst side and second side provide the cooling air holes 86 to injectcooling flow between adjacent blade outer air seals 70.

The location of the cooling holes 86 are described in terms of Cartesiancoordinates indicated by the axes 106 that includes X, Y and Z axeswhich correspond to the axial direction (Y), the circumferentialdirection (X), and the radial direction (Z) as is shown in FIG. 4relative to a point of origin indicated at 84. The locations for thecooling holes 86 correspond to the location where the holes breakthrough the surface of either the first side 74, the second 76 or thegas path surface 72.

The coordinates of the cooling holes are set forth in Table 1 (shownbelow), provide for the circumferential, radial and axial locationrelative to the point of origin 84 on the BOAS 70. Each row in Table 1corresponds with a center line location of an individual hole on one ofthe first side 74, second side 76 and the gas path surface 72. Moreover,each row includes minimum and maximum locations for the each of theholes 86 for each coordinate point. Table 1 includes non-dimensionallocations relative to the point of origin 84. In this example, the pointof origin 84 is disposed on an arc 82 of the forward surface 78. Thepoint of origin 84 is disposed at the center of curvature 80 thatdefines the circumferential radius of the plurality of BOAS around thespecific turbine rotor section.

TABLE 1 Hole X_(min) X_(max) Y_(min) Y_(max) Z_(min) Z_(max) 1 −0.658−0.654 −0.107 −0.107 0.948 0.951 2 −0.516 −0.513 −0.109 −0.109 0.9490.953 3 −0.373 −0.371 −0.101 −0.101 0.951 0.954 4 −0.229 −0.228 −0.107−0.107 0.952 0.955 5 −0.081 −0.081 −0.107 −0.107 0.952 0.955 6 0.0670.067 −0.107 −0.107 0.952 0.955 7 0.215 0.216 −0.107 −0.107 0.952 0.9558 0.362 0.364 −0.107 −0.107 0.951 0.954 9 0.509 0.512 −0.107 −0.1070.949 0.953 10 0.656 0.660 −0.107 −0.107 0.948 0.951 11 −0.664 −0.661−0.271 −0.270 0.947 0.950 12 −0.512 −0.509 −0.277 −0.276 0.949 0.952 13−0.333 −0.331 −0.264 −0.263 0.950 0.953 14 −0.186 −0.185 −0.263 −0.2610.951 0.954 15 −0.044 −0.043 −0.263 −0.261 0.951 0.954 16 0.098 0.099−0.263 −0.261 0.951 0.954 17 0.239 0.241 −0.263 −0.261 0.951 0.954 180.381 0.383 −0.263 −0.261 0.950 0.953 19 0.522 0.524 −0.263 −0.261 0.9480.952 20 0.663 0.666 −0.263 −0.261 0.947 0.950 21 −0.657 −0.654 −0.465−0.463 0.947 0.950 22 −0.487 −0.484 −0.439 −0.437 0.949 0.952 23 −0.335−0.334 −0.439 −0.437 0.950 0.953 24 −0.111 −0.111 −0.424 −0.422 0.9510.954 25 0.044 0.044 −0.424 −0.422 0.951 0.954 26 0.199 0.200 −0.424−0.422 0.951 0.954 27 0.353 0.355 −0.424 −0.422 0.950 0.953 28 0.5080.510 −0.424 −0.422 0.949 0.952 29 0.662 0.665 −0.424 −0.422 0.947 0.95030 −0.657 −0.654 −0.608 −0.605 0.947 0.950 31 −0.502 −0.499 −0.608−0.605 0.949 0.952 32 −0.266 −0.264 −0.570 −0.567 0.951 0.954 33 −0.110−0.110 −0.570 −0.567 0.951 0.954 34 0.045 0.045 −0.570 −0.567 0.9510.954 35 0.200 0.201 −0.570 −0.567 0.951 0.954 36 0.354 0.356 −0.570−0.567 0.950 0.953 37 0.509 0.511 −0.570 −0.567 0.949 0.952 38 0.6630.666 −0.570 −0.567 0.947 0.950 39 −0.722 −0.718 −0.104 −0.104 0.9550.958 40 −0.722 −0.718 −0.255 −0.253 0.955 0.958 41 −0.722 −0.718 −0.404−0.402 0.955 0.958 42 −0.722 −0.718 −0.530 −0.528 0.955 0.958 43 −0.722−0.718 −0.657 −0.653 0.955 0.958 44 −0.722 −0.718 −0.799 −0.795 0.9550.958 45 −0.722 −0.718 −0.941 −0.936 0.955 0.958 46 0.718 0.722 −0.882−0.877 0.955 0.958 47 0.718 0.722 −0.707 −0.704 0.955 0.958 48 0.7180.722 −0.571 −0.568 0.955 0.958 49 0.718 0.722 −0.463 −0.461 0.955 0.95850 0.718 0.722 −0.337 −0.335 0.955 0.958 51 0.718 0.722 −0.219 −0.2180.955 0.958

Each location set forth in Table 1 is where the center line of eachcooling hole 86 breaks through the surface. Additional elements such asadditional cooling holes, protective coatings, and other specificfeatures that would be provided in the BOAS 70 are not described by thecoordinates provided in Table 1.

Manufacturing tolerances are recognized for the fabrication of BOAS 70.Accordingly, the table indicates the tolerance with a minimum andmaximum locations relative to each coordinate point for each location.Moreover, each hole may deviate from a true position with a tolerance ofabout 0.023 inches (0.58 mm) from a center line of the hole. Thespecific tolerance is with regard to the location of each of the holesand generally not scalable although the coordinates provided in thetable are non-dimensional and are therefore scalable relative to thesizes of the blade outer air seal.

Referring to FIG. 7, with continued reference to FIGS. 3, 4, 5 and 6,each of the holes 86 is in communication with at least one passage suchas those indicated at 94 a and 94 b. The passages 94 a and 94 bcommunicate cooling air flow from a supply in communication withopenings on a top side 92 of each BOAS 70. In this example as is shownin FIG. 6, several passages 102 are provided through the BOAS 70 tocommunicate cooling air flow to the film cooling holes 86.

At least one of the passages, for example passage 94 b is disposed at anangle 98 relative to normal 96 to the corresponding surface for thatopening 86. Some of the passages, for example passage 94A may bedisposed normal to the surface through which the opening extends.

Referring to FIGS. 8A and 8B, the holes may include a conical shapecorresponding to a generally oblong opening through the surface. Thisoblong opening provides a direction of air flow once it exits the holes86 to provide the desired flow pattern for cooling air flow. Moreover,the holes 86 may be cylindrically shaped.

Referring to FIG. 9, each of the openings generally include a diameter100 that is measured at least the largest portion of the opening that iswithin a tolerance range of 0.010 to 0.035 inches (0.25-0.89 mm).

Referring back to FIGS. 3, 4 and Table 1, the location of each filmcooling hole 86 is defined according to the table in the circumferential(X), axial (Y) and radial direction relative to the zero point 84. Thelocations are not directional, meaning they indicate a center line ofthe opening regardless of the orientation of the surface through whichit extends. Accordingly, Table 1 defines locations of openings on eachof the first and second sides 74, 76 that are substantially transverseto the gas path surface 72.

Moreover, Table 1 is non-dimensional and scalable and conformance to thedisclosed film cooling hole locations is provided by selecting specificparticular values for the scaling parameters in inches or millimeters.Substantial conformance is based on points representing the cooling holelocations for example, in inches or millimeters as determined byselecting particular values of the scaling parameters.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. A blade outer air seal for a gas turbine enginecomprising: a gas path surface exposed to exhaust gas flow; a first sideextending radially outward from the gas path surface; a second sideextending radially outward from the gas path surface; and a plurality offilm cooling holes disposed on at least one of the gas path surface, thefirst side and the second side, the film cooling holes disposed atlocations described by a set of Cartesian coordinates set forth in Table1, the Cartesian coordinates provided by an axial coordinate, acircumferential coordinate and a radial coordinate relative to a definedpoint of origin.
 2. The blade outer air seal as recited in claim 1,including an axially forward side and an axially aft side, wherein thegas path surface and the forward side define an arc and the point oforigin is defined at the center of curvature of the arc on the forwardside.
 3. The blade outer air seal as recited in claim 2, wherein theforward side and the aft side include features for securement to asupport structure within the turbine section of the gas turbine engine.4. The blade outer air seal as recited in claim 1, the blade outer airseal is one of a plurality of outer air seals disposed circumferentiallyabout a longitudinal axis of the gas turbine engine.
 5. The blade outerair seal as recited in claim 1, wherein each of the film cooling airholes are located within a true position of 0.023 inches (0.58 mm). 6.The blade outer air seal as recited in claim 1, wherein at least some ofthe film cooling air holes comprise one of a conical and cylindricalshape.
 7. The blade outer air seal as recited in claim 1, wherein eachof the film cooling air holes correspond with a passage through thecorresponding surface and at least some of the passages are disposed atan angle different than normal relative to the surface.
 8. The bladeouter air seal as recited in claim 1, wherein a plurality of filmcooling holes have a diameter within a range of 0.010-0.035 inch(0.25-0.89 mm).
 9. A gas turbine engine comprising: a compressor sectiondisposed about an axis; a combustor in fluid communication with thecompressor section; a turbine section in fluid communication with thecombustor, the turbine section includes at least one rotor having aplurality of rotating blades; and a plurality of blade outer air sealscircumferentially surrounding the rotating blades, wherein each of theblade outer air seals includes: a gas path surface exposed to exhaustgas flow; a first side extending radially outward from the gas pathsurface; a second side extending radially outward from the gas pathsurface; and a plurality of film cooling holes disposed on at least oneof the gas path surface, the first side and the second side, the filmcooling holes disposed at locations described by a set of Cartesiancoordinates set forth in Table 1, the Cartesian coordinates provided byan axial coordinate, a circumferential coordinate and a radialcoordinate relative to a zero-coordinate.
 10. The gas turbine engine asrecited in claim 9, wherein each of the blade outer air seals includesan axially forward side and an axially aft side, wherein the gas pathsurface and the forward side define an arc and the zero-coordinate isdefined at the center of curvature of the arc on the forward side. 11.The gas turbine engine as recited in claim 10, wherein the forward sideand the aft side include features for securement to a support structurewithin the turbine section of the gas turbine engine.
 12. The gasturbine engine as recited in claim 9, wherein each of the plurality offilm cooling air holes are located within a true position of 0.023inches (0.58 mm).
 13. The gas turbine engine as recited in claim 9,wherein at least some of the plurality of film cooling air holescomprise one of a conical and cylindrical shape.
 14. The gas turbineengine as recited in claim 9, wherein each of the plurality of filmcooling air holes are in communication with a corresponding plurality ofpassages and at least some of the passages are disposed at an angledifferent than normal relative to the surface.
 15. The gas turbineengine as recited in claim 9, wherein a plurality of film cooling holeshave a diameter within a range of 0.010-0.035 inch (0.25-0.89 mm).